JPH08337200A - Solar search method of triaxial stabilizing satellite and triaxial stabilizing satellite - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a sun sensor with a view field having a gap to be utilized only in the case where a measurement axis is oriented as desired relatively to the view field by generating a moment which is not parallel to the direction of the measurement axis by a flywheel device to be additively used. SOLUTION: A control device is provided with a rotation speed gyro, and a rotation moment generating device which is not a flywheel device, and the rotation moment generating device is controlled based on a measurement signal of the rotation speed gyro. At this time, control τ=-kGGTω is used. Where, τis an inertia moment, k is a coefficient, G is a direction vector, and ω is a rotation vector of a satellite. Two reaction wheels RWX and RWZ of which rotation axes respectively coincide with X-axis and Y-axis, and two spin wheels FMW1, FMW2 of which rotation axes form the same angles θ to negative Y-axis with the opposite signs to each other, and which exist on the same plane with the axes are provided. A corresponding surface forms an angle ηto negative X-axis.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a sun search method for a three-axis stabilizing satellite, and this three-axis stabilizing satellite.

[0002]

2. Description of the Related Art According to the specification of WO 93/04923 A1, the measuring device used for attitude control of a three-axis stabilizing satellite, and the sun from any initial attitude of the satellite where the sun is not yet present in the field of view of the sun sensor Methods of capturing are known. This measuring device includes a sun sensor that detects the sun direction with respect to a coordinate system fixed to the satellite, and a rotation speed gyro that is measured on a single axis. The field of view or measurement range of the sun sensor should cover a total of 2π total angles within a preselected plane of the coordinate system, and only within a limited angular range <π / 2 in each of two directions perpendicular to said plane. Cover. In this case, the allowable direction of the measuring axis of the rotation speed gyro is restricted,
This limit depends on the width of the field of view of the sun sensor perpendicular to the preselected plane. The angle α which defines the field of view of the sun sensor in the preselected plane and perpendicular to this plane, mentioned there.
_The conditions for the allowable range of ₁ or α ₂ and the angle β formed between the measuring axis of the rotational speed gyro and the preselected surface are defined as follows. That is, 0 ≤ α ₁ ≤ 2π (1a) -α _2max ≤α ₂ ≤α _2max (1b) _│β│ ≥ (π / 2) -α _2max (1c) _{Preliminarily} captured by the field of view of the solar sensor device The selected plane is preferably the XZ plane of the coordinate system fixed to the satellite. In the case of a geostationary satellite that is orientated as specified, the X axis indicates the roll axis in the orbit direction, and the Z axis indicates the earth's roll axis. With the yaw axis facing the center of gravity, the Y axis is the pitch axis perpendicular to both other axes.

However, due to other operational requirements, it may not be possible to satisfy the above-mentioned conditions regarding the setting of the measuring axis of the rotation speed gyro. this is,
For example, the requirement applies for β _max ≤ (π / 2) − α _2max (2). Furthermore, the above requirement that the field of view of the sun sensor device should have a full angle of 2π in a preselected plane is not always permitted and is a non-optimal limitation from a cost standpoint. This is because it is necessary to use a corresponding number of sun sensors to cover the entire 2π angular range. Furthermore, when one or more sun sensors fail, there is a gap in the field of view, which makes the measurement and evaluation methods provided, and especially the method provided for capturing the sun, also infeasible. To The same applies if the surrounding field of view is limited by a protruding antenna or other device mounted on the satellite.

[0004]

SUMMARY OF THE INVENTION The object of the invention is to work even if the limits defined above with regard to the field of view of the sun sensor and the setting of the measuring axis of the rotational speed gyro cannot be observed, and the field of view of the sun sensor device is preset. It is an object of the present invention to provide a solar probing method of the kind mentioned at the outset for a triaxial stabilizing satellite which can be used even if it has a gap in the selected plane and the measuring axis is arbitrarily oriented with respect to the field of view. It is a further object of the invention to present a three-axis stabilized satellite which, due to the device configuration, can successfully carry out the method of the invention for searching the sun.

[0005]

According to the invention, the above objects are separated from one another by a partial section or gap of a preselected plane of an orthogonal three-axis coordinate system whose field of view is fixed to the satellite. A sun sensor device that covers a plurality of sub-sections and a range perpendicular to the plane, a rotation axis gyro that has a measurement axis with an arbitrary orientation, and measures a single axis, and a rotation moment around all three coordinate axes. A rotation moment generator that is not a flywheel device to generate, and a control device that outputs a control signal of the rotation moment generator based on a measurement signal of a rotation speed gyro to generate a control moment, and a control law of the following form, τ = −k GGT ^T ω , where τ is the control moment, k is the amplification factor, and G
Is the direction vector of the measurement axis of | G | = 1, G ^T is the transposed vector with respect to G , and ω is the rotational velocity vector of the satellite,
A method for searching the sun of a three-axis stabilizing satellite is solved by generating a rotational moment (rotational moment vector H 1 ) which is not parallel to the direction of the measurement axis by a Philawheel device to be additionally used.

Further, according to the present invention, the above-mentioned problem is that the field of view in a preselected plane of the three-axis orthogonal coordinate system fixed to the satellite covers one section or a plurality of sections separated from each other by a gap. Set the sensor device and measurement axis in any direction (|
Rotation speed gyro that measures with a single axis directed to the direction vector G 1) where G | = 1, rotation moment generator that generates rotation moments around all three axes, and rotation moment generation based on the measurement signal of the rotation speed gyro A three-axis stabilizing satellite with an adjusting device for outputting the adjusting signal of the device is solved by further providing a flywheel device for generating a rotational moment component around all three coordinate axes.

Other advantageous configurations according to the invention are described in the dependent claims.

[0008]

Accordingly, a flywheel device is also used. Rotational pulses are generated with the aid of this device. The accessory rotation vector H of this pulse is not oriented parallel to the direction of the measurement axis of the rotation speed gyro. in this case,
(As already known from WO 93/04923 A1 specification)
It is assumed that a constant control rule is used that nulls the rotational speed component oriented parallel to the measuring axis of the rotational speed gyro. Generating an additional rotation pulse will only leave at most satellite rotation about the rotation pulse axis in the control process. This rotation is very often enough that the sun is finally in the field of view of the sun sensor device.

The dependent claims 2 to 8 show advantageous configurations of two embodiments of the method according to the invention. These embodiments relate on the one hand to the case where the field of view of the sun sensor device covers at least an angle of π in a preselected plane without a gap, and on the other hand the field of view has a partial area interrupted by a gap, which is small. This is relevant when the other subsection covers the angle π. Claim 9
Defines the configuration of a three-axis stabilizing satellite necessary for the successful implementation of the method according to the invention. This satellite does not have the restrictions mentioned at the beginning with respect to the orientation of the sun sensor device and the measuring axis of the rotational speed gyro that measures in a single axis, i.e. the entire field of view of the sun sensor device is a preselected plane. The only measuring axis of the rotational speed gyro that is not inside can be oriented arbitrarily. An important feature is the addition of a flywheel device that can generate rotational moment components around all three coordinate axes. When the field of view of the sun sensor is strongly restricted, no satellite of this kind has ever been designed with a single rotational speed gyro that is measured in a single axis to measure the rotational speed. Because it is measured in three axes, and especially in the case of redundant designs, if there is a correspondingly expensive and prone to failure speed gyro, the required initial sun search or constant calculation of the required sun search will occur. This is because the problems that should be put into can only be overcome so far. Here, for the first time, a satellite in time with a single-axis rotational speed gyro for an important operation of the sun search is introduced, making this operation possible using existing flywheel equipment.

Dependent claims 10 to 12 disclose other advantageous configurations of this satellite with respect to the configuration of the flywheel arrangement.

[0011]

Embodiments of the present invention will now be described in more detail with reference to the accompanying drawings. FIG. 1 shows the required field of view and extent of the sun sensor device, which is known from WO 93/04923, and of the measuring axis of a gyro measured in a single axis in a coordinate system XYZ fixed to the satellite. The allowed orientations (direction vector G 1 ) are shown. This sun sensor device has a full field of view (0 ≤ α ₁ ≤ 2π) in the XZ plane, and a field of view in two directions perpendicular to this direction is limited to ± α 2 _max , so the positive and negative Y-axis There is a double cone shaped area around which the sun sensor device does not capture. Only the hatched area is allowed for the direction vector G of the rotation speed gyro. These areas are determined by the following requirements. That is, | β | ≧ (π / 2) −α _2max (1c) where β is the angle between the XZ plane and the direction vector G. However, this requirement is not always sustained. In other words, it is not maintained due to the demands determined by other operational hardness. In this case, the solar search method presented in WO 93/04923 can no longer be used.

[0012] This is sufficient for full field of view in the case of normal functioning if some of the full field of view, for example the individual sun sensors covering 1/3, have failed, or for cost reasons. The same applies if a number of individual sun sensors need to be omitted. The arrangement that disturbs the above conditions covers, for example, an angle in which the field of view of the sun sensor in the XZ plane of the coordinate system fixed to the satellite is larger than π, and the direction vector G of the rotation speed gyro measured by a single axis is also XZ. In the plane, especially in the center of this field of view. This type of arrangement is shown in FIG. The field of view in the XZ plane is shown hatched in FIG. 2 and is centered around the direction vector G of the measuring axis of the rotational speed gyro. In particular,
The direction vector G also matches the X axis. The field of view has at least two vectors ± e _R. These vectors are

[0013]

[Outside 2]

Is defined as follows, and e _Y is a unit vector in the Y-axis direction. Naturally, the field of view of the sun sensor device thus defined is perpendicular to the XZ plane, for example ± 30
Has an extension of °. In that case, this value will be somewhat smaller at the edges. A three-axis stabilized satellite with this kind of special arrangement of the field of view of the sun sensor and the measuring axis of the rotational speed gyro normally uses a rotational moment generator for attitude control. The device generally consists of three pairs of attitude control nozzles, each of which produces positive and negative rotational moments about a coordinate axis. Instead of the attitude control nozzles, a system of electromagnetic coils may be present if there are other devices that generate rotational moments, for example an external magnetic field.

According to the present invention, there is further provided a rotation moment generating device capable of generating a rotation moment component in all three coordinate directions. In this case, the important thing is the flywheel device. In the simplest case, this device consists of only one flywheel which can be tilted in each of the desired directions, but it generally has at least three individual flywheels, the rotary axes of which are all It's not in one plane. By appropriately adjusting the rotational speed and the rotational direction of these flywheels, it is possible to generate a combined rotational moment (rotational moment vector H 1 ) of a desired value and a desired direction. A preferred embodiment of this type of filly wheel device is shown in FIG. This figure 3
Include two reaction wheels RWX and RWZ whose rotation axes coincide with the X-axis and the Z-axis, respectively, and a rotation axis which forms the same angle θ as the negative Y-axis but with opposite signs. Two spin wheels are shown which are coplanar with the axis. The corresponding surface forms an angle η with the negative X axis. A typical value for these angles is η = 45 ° θ = 10 °.

This type of arrangement of flywheels can be defined as follows, independent of the coordinate axes. That is, there are two spin wheels having an acute-angled rotation axis and two reaction wheels arranged so as to be orthogonal to each other with respect to the rotation axis, and both rotation axis planes stretched by the two rotation axes are orthogonal to each other. , The angle halved by each of the two pairs of wheels lies in the plane perpendicular to the two axes of rotation. Spin wheels are generally designed to provide and maintain a large rotational moment in one direction, but reaction wheels produce relatively small rotational moments in two rotational directions. The flywheel device of FIG. 3 is preferably designed to impart a rotational moment vector H in the direction of the negative Y axis. In that case, some kind of misalignment in the other two coordinate directions can be easily adjusted. This device is designed with redundancy so that even if any one of the four flywheels fails, the rotational moment vector H can be generated in the substantially negative Y-axis direction or in any other direction. You can

Since a sun search is performed using this type of sun sensor device, the rotation speed is known when it is known with a certain accuracy.
First, it is controlled and attenuated by a known method. This is the case, for example, after disconnecting the satellite from the carrier rocket. To this end, the rotational moment generator τ described above, that is, the attitude control nozzle, for example, gives a rotational moment τ in the opposite direction to the rotational speed vector ω . This moment acts during the following time interval Δt Δt = | Iω | / | τ | (4). In this case, I is the satellite inertia tensor. During this process, the flywheel of the flywheel device is stopped. After that, a small unknown residual rotational speed ω _R generally remains. This is reduced in the following steps.

In this respect, the invention is introduced, but it can also be used without the method steps already described. Using the flywheel device, a constant rotational moment vector H
, Which in turn requires only that the vector be oriented so that it is not parallel to the direction vector G of the measuring axis of the rotational speed gyro. As already mentioned, another adjusting device is required, which produces an adjusting moment, so that it rotates on the basis of the rotational speed gyro measurement signal as if it were present in the attitude control device of each satellite. The adjustment signal of the moment generator is output.
Therefore, the system equations arise. This ignores the Euler term,

[0019]

[Outside 3]

[0020]

[Outside 4]

The second term on the right side of (1) corresponds to the control law known from WO 93/04923 A1. This control law attenuates the component of the rotation speed vector ω parallel to the direction vector G to zero. The first term, which depends on the rotational moment vector H , is due to the fact that the rotational speed component oriented perpendicular to H is linked to the measuring axis G. In this case, the component of the rotation speed vector ω that is perpendicular to the direction vector G and perpendicular to the rotation moment vector H always carries away the rotation energy and inputs it to the rotation shaft. Therefore, this component is attenuated by the second term. This is done using an attitude control system, for example using an attitude control nozzle that produces an adjustment moment τ .

Optimal damping of the component of the rotational velocity vector ω that is perpendicular to the direction vector G is achieved when the rotational moment H is perpendicular to G. At that time, as a result of the control process, only a slight rotation around H remains,
This can be zero. In general, the above system equations can also be formulated using the system matrix A as That is,

[0023]

[Outside 5]

It is The selection of the vectors H and G determines the eigenvalues and damping characteristics of the system matrix A. Let these vectors be G
And H are chosen such that they deviate from the rule of vertically overlapping each other, adjusting the rotational velocity component around all three coordinate axes to zero, but with strong oscillatory transients. This is usually undesirable. A particularly important practical case is that G and H lie in the XZ plane and are oriented perpendicular to each other, as also shown in FIG. With this arrangement, the rotational velocity component along the direction axis G and the Y axis of the coordinate system fixed to the satellite is completely damped, so that only the rotation around the rotational moment vector H remains, as already explained above. .

When the rotational moment vector H is oriented as shown in FIG. 2, that is, at the edge of the field of view of the sun sensor device which covers at least the angle π as a whole and also contains the vector H , the residual satellites around H The rotation gives the capture of the sun sooner or later. However, when this residual velocity is very small or zero, another exploration action begins. This probing action activates a rotation about the axis e _R , using a rotational moment generator, ie, an attitude control nozzle, which rotation is in the field of view of the sun sensor device at XZ,
In other words, this is wide in the edge part so that it also applies to the vector e _R. As a preferred axis, e _R = H
You can choose to. This is the case for optimal consumption. The time duration of the required rotational moment pulse is predicted based on the residual velocity and remains due to the damping controlled using the control law above.

If the field of view of the sun sensor device in the XZ plane consists of one or more subsections separated from one another by a gap, the smaller subsections covering at least the relevant angular range of π Method step, ie, controlled damping using the control law above, and other final exploration behavior. It commands the rotation of the satellite around one axis of rotation, or the continuous rotation around multiple axes of rotation, with the adjustment of the The final step is to make sure to soak. This kind of behavior is described, for example, in German Patent No.
7 49 868 C3 specification or document "The Attitude a
nd Orbit Control Subsystem of the TV-SAT / TDF1 Spac
ecraft "by H. Bittner et al., IFAC-Symposium on Au
tomaticControl in Space, 1982, pp. 83-102. However, this method presupposes a rotational speed gyro that measures on three axes or an equivalent measuring device.
In this case, in order to take into account the fact that only a single rotational speed gyro is used, the unknown component of the rotational speed vector ω , which is oriented perpendicular to the direction vector G of the measuring axis, is predicted. Use the body. This has, for example, the structure shown in FIG.

The actual satellite dynamics are shown in the dashed box and the equation

[0028]

[Outside 6]

It becomes Where τ is the adjustment moment introduced into the satellite. The output amount y is measured by the rotation speed gyro 2 which is measured on a single axis. This amount is proportional to the rotation speed component perpendicular to the direction vector G of the measurement axis of the rotation speed gyro. In other words, it is y~ G ^T ω (7).

The signals corresponding to the adjustment moment τ and the output signal y also reproduce the dynamic characteristics of the satellite in detail.

[0031]

[Outside 7]

And its time derivative is formed. The quantity s usually represents a Laplace operator or a differential operator. In addition, the satellite inertial matrix I is used. In the arrangement shown, the observation object 3 can be realized as a hard-wired connection or a pure algorithm of an in-vehicle computer of the satellite. Addition point 6
The difference value formed by means of selectively passes through the amplifier 7, in which the vector amplification factor 1 of l ^T = (1 _1, 1 _2, 1 ₃ ) is used. Formed by observing bodies

[0033]

[Outside 8]

It is introduced into the adjusting device 4. This adjusting device has the following control law,

[0035]

[Outside 9]

In order to realize the above, the following adjustment signal of the rotational moment generator 5 is generated. this

[0037]

[Outside 10]

This is a vector for the rotation speed performed at the time of explicit search. This vector must be determined by taking into account the size and orientation of the individual sun sensor's field of view to ensure that the sun will eventually immerse in one of the sub-intervals, and in some cases, constantly in multiple consecutive steps. Need to change. In this case, it should be noted that the flywheel device according to the present invention can also provide the rotational moment vector H during this final exploration motion, which vector is given by the observable matrix Q _B attached to the observing body. It has a rank of 3. This matrix is well known

[0039]

[Outside 11]

Is given by If the three column vectors of this matrix are linearly independent of each other, then there is observability. This adjusts the rotation moment vector H and 45 ° between H and G.
Angle is generated. In this case, all three axes of motion are tied together. The prerequisite for using this exploration method is that the satellite is still spinning at a rotational speed small enough to satisfy the following inequality | I ω | << | H | (10) after the end of controlled damping. is there. The following exercise methods

[0041]

[Outside 12]

[0042]

[Outside 13]

Then, the observation equation described above occurs.
However, the above preconditions are always given after the controlled damping according to the invention. If the satellite has a spherical mass distribution

[0044]

[Outside 14]

Finally, FIG. 5 shows in a related diagram the individual substeps of the search strategy for the sun search according to the invention.

[0046]

As described above, when the solar exploration method of the present invention for a three-axis stabilizing satellite cannot observe the restrictions defined above regarding the setting of the field of view of the sun sensor and the measurement axis of the rotation speed gyro. However, it can be used even when the field of view of the sun sensor device has a gap in a preselected plane and the measuring axis is oriented arbitrarily with respect to the field of view.

[Brief description of drawings]

FIG. 1 is a drawing showing the position and extent of the field of view of the satellite sun sensor device and the allowable arrangement of the measurement axis of the gyro for single axis measurement

FIG. 2 is a drawing when the field of view of the sun sensor is in the XZ plane of the coordinate system fixed to the satellite.

FIG. 3 is a layout diagram of a flywheel device,

FIG. 4 is a block circuit diagram showing a control system of an observation object,

FIG. 5 is a control flowchart.

[Explanation of symbols]

1 Satellite dynamics measurement unit 2 Rotational speed gyro 3 Observer 3 Adjustment device 5 Rotational moment generation device 6 Addition point 7 Amplifier RWX, RWZ Reaction wheel FMW1, FMW2 Spin wheel H Rotational moment vector G Direction vector of measurement axis

─────────────────────────────────────────────────── ─── Continuation of the front page (51) Int.Cl. ⁶ Identification code Internal reference number FI technical display location // G05D 1/08 G05D 1/08 A (72) Inventor Christie Lohier, Federal Republic of Germany, 80796 Miyunchen, Clemensstraße, 70 (72) Inventor Walter Huichter, Federal Republic of Germany, 81827 Miyunchen, Kestrenstraße, 20

Claims (12)

[Claims] 1. An orthogonal triaxial seat whose field of view is fixed to the satellite.
In one subsection or gap of a preselected face of the standard
A plurality of subsections that are separated from each other and the surface
A sun sensor device that covers a certain straight range and a measurement in any direction.
Rotation speed gyro that has a fixed axis and measures with a single axis, and three
To generate rotational moments around all the coordinate axes of
A rotation moment generator that is not a rye wheel device,
A control signal is generated for the rotation speed gyro to generate a control moment.
Output the control signal of the rotation moment generator based on the
Control device of the following form,τ= -KGG
^T ωUse hereτIs the control moment, k is the amplification factor
number,GIs |GThe direction vector of the measurement axis with | = 1,G ^TButG
The transpose vector for,ωIs the satellite rotation speed vector
An additional method in the sun search method of a three-axis stabilizing satellite
The direction of the measuring axis depends on the Philiwheel device that should be used for
Moment not parallel to (rotation moment vector
H) Is generated. 2. Using a sun sensor device whose field of view covers at least an angular range of π in a preselected plane with no space, the rotational moment H is oriented perpendicular to the direction vector G of the measuring axis. The method of claim 1, wherein: 3. The orientation of the rotation moment vector H is adjusted so that the rotation moment vectors H and −H are also within the field of view of the sun sensor device starting from the coordinate origin of a coordinate system fixed to the satellite. The method of claim 2 characterized. 4. If there is no or only a slight rotation of the satellite around the direction of the rotation moment vector H using the control law, command rotation around the vector e _R , which rotation is the sun. Vector in the field of view of the sensor device- e
The method of claim 3, wherein the method is in a preselected plane similar to _R. 5. The vector e _R is a rotational moment vector.
The method of claim 4, wherein the method is collinear with H. 6. A solar sensor device is used, wherein the preselected field of view does not have a partial section that covers at least the angular range of π without a gap, and a matrix The method of claim 1, wherein the rotational moment vector H is oriented such that has a rank of 3. 7. The rotational moment vector H is the same as the measuring axis G.
The method according to claim 6, wherein the angle is 45 °. 8. Taking into account the azimuth and size of a partial section of the field of view of the sun sensor device, the satellite may be rotated once around one axis of rotation, or sequentially rotated around a plurality of axes of rotation. Command to ensure that the sun will eventually enter one of the subintervals, in order to predict an unknown component of the rotational velocity vector ω that is oriented perpendicular to the measurement axis. 8. The method according to claim 6 or 7, characterized in that 9. A sun sensor device that covers a single section or a plurality of sections separated from each other by a gap in a plane selected in advance in a three-axis orthogonal system fixed to a satellite, and an arbitrary measuring axis. direction and speed gyros for measuring a single axis oriented in (| | G = 1 direction vector G is a), the torque generating device that generates a rotational moment around three axes all, based on the measurement signal of the rotational speed gyro A three-axis stabilizing satellite equipped with an adjusting device for outputting an adjusting signal of a rotating moment generating device, further comprising a flywheel device for generating a rotating moment component around all three coordinate axes. . 10. The satellite according to claim 9, wherein the flywheel device has at least three flywheels which do not have all rotation axes in one plane as a whole. 11. Satellite according to claim 10, characterized in that there are at least one reaction wheel and at least one spin wheel each. 12. There are two spin wheels each having an acute-angled rotation axis and two reaction wheels arranged so as to be orthogonal to each other with respect to the rotation axis, and both rotation axis surfaces stretched by the two rotation axes. Are orthogonal and the bisector of the angle spanned by the two pairs of wheels lies in a plane oriented orthogonal to both planes of rotation.
The satellite according to 1.